Geared gas turbine engine with reduced oil tank size

ABSTRACT

A gas turbine engine comprises a fan drive turbine for driving a gear reduction, which drives a fan rotor. A lubrication system supplies oil to the gear reduction. An oil tank is relatively small. The lubrication system operates to allow oil to remain in the oil tank for a dwell time of less than or equal to five seconds. A method of designing a gas turbine engine is also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.14/595,255 filed Jan. 13, 2015, which claims priority to U.S.Provisional Patent Application No. 61/929,174, filed Jan. 20, 2014.

BACKGROUND OF THE INVENTION

This application relates to a gas turbine engine having a gear reductiondriving a fan wherein an oil tank is of reduced size.

Gas turbine engines are known and, typically, include a fan deliveringair into a bypass duct as propulsion air. The fan also delivers air intoa core engine where it passes to a compressor. The air is compressed inthe compressor and delivered downstream into a combustion section whereit is mixed with fuel and ignited. Products of this combustion passdownstream over turbine rotors driving them to rotate.

Historically, the fan rotor and a fan drive turbine rotor have beendriven at the same speed. This placed a restriction on the desirablespeed of both the fan and the fan drive turbine.

More recently, it has been proposed to provide a gear reduction betweenthe fan drive turbine and the fan rotor.

The gear reduction is a source of increased heat loss. As an example, ageared turbofan engine creates about twice as much heat loss as anon-geared turbofan engine. In addition, the weight of the engineincreases due to the weight of the gear reduction.

It has typically been the case that a designer of a gas turbine enginesizes an oil tank such that the oil can sit in the oil tank long enoughto de-aerate. On a normal turbofan engine, this had been approximatelyat least ten seconds.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine comprises a fan driveturbine for driving a gear reduction, which drives a fan rotor. Alubrication system supplies oil to the gear reduction. An oil tank isrelatively small. The lubrication system operates to allow oil to remainin the oil tank for a dwell time of less than or equal to five seconds.

In another embodiment according to the previous embodiment, the dwelltime is less than or equal to 3.0 seconds.

In another embodiment according to any of the previous embodiments, thegear reduction includes a sun gear for driving intermediate gears. Thereare oil baffles located circumferentially between the intermediategears.

In another embodiment according to any of the previous embodiments, anoil capture gutter surrounds the gear reduction.

In another embodiment according to any of the previous embodiments, thefan rotor delivers air into a bypass duct as propulsion air and into acore engine where it passes into a compressor section. A bypass ratio isdefined as the ratio of air delivered into the bypass duct compared tothe volume of air delivered into the core engine. The bypass ratio isgreater than or equal to about 6.0.

In another embodiment according to any of the previous embodiments, thebypass ratio is greater than or equal to about 10.0.

In another embodiment according to any of the previous embodiments, agear ratio of the gear reduction is greater than or equal to about2.3:1.

In another embodiment according to any of the previous embodiments, theoil tank may hold greater than or equal to 25 and less than or equal to35 quarts of oil.

In another embodiment according to any of the previous embodiments, theengine is rated greater than or equal to 15,000 and less than or equalto 35,000 lbs in rated thrust at take-off.

In another embodiment according to any of the previous embodiments, theoil tank holds greater than or equal to 35 and less than or equal to 50quarts of oil.

In another embodiment according to any of the previous embodiments, theoil tank is associated with an engine having greater than or equal to35,000 and less than or equal to 100,000 lbs in rated thrust attake-off.

In another embodiment according to any of the previous embodiments, thegear reduction includes a sun gear for driving intermediate gears. Thereare oil baffles located circumferentially between the intermediategears.

In another embodiment according to any of the previous embodiments, anoil capture gutter surrounds the gear reduction.

In another embodiment according to any of the previous embodiments, theoil tank may hold greater than or equal to 25 and less than or equal to35 quarts of oil.

In another embodiment according to any of the previous embodiments, theengine is rated greater than or equal to 15,000 and less than or equalto 35,000 lbs in rated thrust at take-off.

In another embodiment according to any of the previous embodiments, theoil tank holds greater than or equal to 35 and less than or equal to 50quarts of oil.

In another embodiment according to any of the previous embodiments, theoil tank is associated with an engine having greater than or equal to35,000 and less than or equal to 100,000 lbs in rated thrust attake-off.

In another embodiment according to any of the previous embodiments, theoil tank may hold greater than or equal to 25 and less than or equal to35 quarts of oil. The engine is rated greater than or equal to 15,000and less than or equal to 35,000 lbs in rated thrust at take-off.

In another embodiment according to any of the previous embodiments, theoil tank holds greater than or equal to 35 and less than or equal to 50quarts of oil. The oil tank is associated with an engine having greaterthan or equal to 35,000 and less than or equal to 100,000 lbs in ratedthrust at take-off.

In another embodiment according to any of the previous embodiments, anoil capture gutter surrounds the gear reduction.

In another featured embodiment, a method of designing a gas turbineengine comprises the steps of providing a fan drive turbine for drivinga gear reduction. The gear reduction drives a fan rotor. A lubricationsystem is provided to supply oil to the gear reduction. An oil tank isrelatively small. The lubrication system operates to allow oil to remainin the oil tank for a dwell time of less than or equal to five seconds.

In another embodiment according to the previous embodiment, the dwelltime is less than or equal to 3.0 seconds.

In another embodiment according to any of the previous embodiments, thegear reduction includes a sun gear for driving intermediate gears. Thereare oil baffles located circumferentially between the intermediategears.

In another embodiment according to any of the previous embodiments, anoil capture gutter around the gear reduction is provided.

In another embodiment according to any of the previous embodiments, thefan rotor delivers air into a bypass duct as propulsion air and into acore engine where it passes into a compressor section. A bypass ratio isdefined as the ratio of air delivered into the bypass duct compared tothe volume of air delivered into the core engine. The bypass ratio isgreater than or equal to about 10.0.

In another embodiment according to any of the previous embodiments, agear ratio of the gear reduction is greater than or equal to about2.3:1.

In another embodiment according to any of the previous embodiments, theoil tank may hold greater than or equal to 25 and less than or equal to35 quarts of oil.

In another embodiment according to any of the previous embodiments, theengine is rated greater than or equal to 15,000 and less than or equalto 35,000 lbs in rated thrust at take-off.

In another embodiment according to any of the previous embodiments, theoil tank holds greater than or equal to 35 and less than or equal to 50quarts of oil.

In another embodiment according to any of the previous embodiments, theoil tank is associated with an engine having greater than or equal to35,000 and less than or equal to 100,000 lbs in rated thrust attake-off.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of a gas turbine engine.

FIG. 2 shows a portion of a gear reduction.

FIG. 3 shows another portion of a gear reduction.

FIG. 4 shows a lubrication system.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than or equalto about six (6), with an example embodiment being greater than aboutten (10), the geared architecture 48 is an epicyclic gear train, such asa planetary gear system or other gear system, with a gear reductionratio of greater than about 2.3 and the low pressure turbine 46 has apressure ratio that is greater than about five. In one disclosedembodiment, the engine 20 bypass ratio is greater than or equal to aboutten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor 44, and the low pressure turbine 46 has apressure ratio that is greater than about five 5:1. Low pressure turbine46 pressure ratio is pressure measured prior to inlet of low pressureturbine 46 as related to the pressure at the outlet of the low pressureturbine 46 prior to an exhaust nozzle. The geared architecture 48 may bean epicycle gear train, such as a planetary gear system or other gearsystem, with a gear reduction ratio of greater than about 2.3:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFCT’)” is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed”-is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram °R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

As shown in FIG. 2, a flexible shaft 99, which is driven by the turbine46, drives a sun gear 101 which, in turn, engages and drivesintermediate gears 102. In some embodiments, the intermediate gears 102may be planet gears of a planetary epicyclic gear system. In otherembodiments, the intermediate gears 102 may be star gears of a starepicyclic gear system. The intermediate gears 102 engage and drive aring gear 103 to, in turn, drive an output shaft 106, which then drivesthe fan rotor 42. In other embodiments, a planetary gear carrier (notshown) driven by planetary gears may drive the fan shaft. Lubricant issupplied to a journal pin 108, to the intermediate gears 102 and toother locations within the gear reduction 48.

FIG. 3 shows baffles 100 which are placed circumferentially betweenadjacent planet gears 102.

A gutter 104 surrounds the gear reduction 48 and captures oil that hasleft the gear reduction. Oil from the gear reduction 48 is returned to apump 72 (See FIG. 4) or a tank 90 as shown schematically in FIG. 4. Asshown, a lubricant system 70 includes the gear reduction 48 which may bestructured as shown in FIGS. 2 and 3. Notably, complete details of theoperation of the baffle, the gutter and the other portions of the gearreduction may be as disclosed in U.S. Pat. No. 6,223,616, the disclosureof which with regard to the operation of the gear reduction isincorporated by reference.

Oil flows from an oil pump 72 to a filter 74 through a pressure reliefvalve 76 to an air/oil cooler 78 and then to a fuel/oil cooler 80. Theoil may pass through an oil pressure trim orifice 82 and back to thetank 90. Alternatively, the oil may pass through a strainer 84 and thento the gear reduction 48. Oil returning from the gear reduction and, inparticular, from the gutter, may pass back directly to the pump 72 or tothe tank 90. This is a simplification of the overall lubricant systemand, as appreciated, there may be other components.

Applicant has recognized that by utilizing baffles 100 and a gutter 104on the gear reduction 48, which may be generally as disclosed in theabove-mentioned U.S. Patent, the oil need not sit in the oil tank forten seconds in order to de-aerate. Thus, the size of the tank 90 may bemade much smaller.

Applicant has discovered that oil is de-aerated by the baffles 100 andgutter system and that a dwell time in the oil tank to remove airbubbles may be less than five seconds. More preferably, it may be lessthan or equal to about 3.0 seconds. This allows the use of oil tank 90to be of a size roughly equivalent to the size utilized in priornon-geared gas turbine engines.

As an example, an oil tank that holds 25 to 35 quarts of oil may beutilized on a geared gas turbine engine with 15,000 to 35,000 lbs inrated thrust at take-off. Further, an oil tank may be 35 quarts to 50quarts of oil for an engine with 35,000 to 100,000 lbs in rated thrustat take-off.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a fan drive turbine driving a gearreduction, said gear reduction driving a fan rotor, and said gearreduction including a sun gear driving intermediate gears; a lubricationsystem supplying oil to said gear reduction and comprising an oil tank,wherein the oil remains in the oil tank for a dwell time of less than orequal to five seconds during operation; an oil capture guttersurrounding said gear reduction; and a gear ratio of said gear reductiongreater than or equal to 2.3:1.
 2. The gas turbine engine as set forthin claim 1, wherein said dwell time is less than or equal to 3.0 secondsduring operation.
 3. The gas turbine engine as set forth in claim 2,further comprising oil baffles located circumferentially between saidintermediate gears.
 4. The gas turbine engine as set forth in claim 3,wherein said fan rotor delivers air into a bypass duct as propulsion airand into a core engine where it passes into a compressor section, with abypass ratio defined as the ratio of air delivered into said bypass ductcompared to the volume of air delivered into said core engine, saidbypass ratio being greater than or equal to 10.0.
 5. The gas turbineengine as set forth in claim 4, wherein said oil tank holds greater thanor equal to 25 and less than or equal to 35 quarts of oil.
 6. The gasturbine engine as set forth in claim 4, wherein said oil tank holdsgreater than or equal to 35 and less than or equal to 50 quarts of oil.7. The gas turbine engine as set forth in claim 6, wherein said oil tankis associated with an engine having greater than or equal to 35,000 andless than or equal to 100,000 lbs in rated thrust at take-off.
 8. Thegas turbine engine as set forth in claim 7, wherein the fan driveturbine includes an inlet, an outlet, and a fan drive turbine pressureratio greater than 5, wherein the fan drive turbine pressure ratio is aratio of a pressure measured prior to the inlet as related to a pressureat the outlet prior to any exhaust nozzle, and the fan rotor includes aplurality of fan blades, a fan pressure ratio across the fan blades ofless than 1.45, measured across the fan blades alone.
 9. The gas turbineengine as set forth in claim 1, wherein said fan rotor delivers air intoa bypass duct as propulsion air and into a core engine where it passesinto a compressor section, with a bypass ratio defined as the ratio ofair delivered into said bypass duct compared to the volume of airdelivered into said core engine with said bypass ratio being greaterthan or equal to 10.0.
 10. The gas turbine engine as set forth in claim9, wherein said oil tank holds greater than or equal to 25 and less thanor equal to 35 quarts of oil.
 11. The gas turbine engine as set forth inclaim 9, wherein said oil tank holds greater than or equal to 35 andless than or equal to 50 quarts of oil.
 12. A gas turbine enginecomprising: a fan drive turbine driving a gear reduction, said gearreduction driving a fan rotor, and said gear reduction including a sungear driving intermediate gears; a lubrication system supplying oil tosaid gear reduction and comprising an oil tank, wherein the oil remainsin the oil tank for a dwell time; a gear ratio of said gear reductiongreater than or equal to 2.3:1; and wherein said dwell time is less thanor equal to 3.0 seconds during operation.
 13. The gas turbine engine asset forth in claim 12, further comprising oil baffles locatedcircumferentially between said intermediate gears.
 14. The gas turbineengine as set forth in claim 13, further comprising an oil capturegutter surrounding said gear reduction.
 15. The gas turbine engine asset forth in claim 14, wherein said oil tank holds greater than or equalto 25 and less than or equal to 35 quarts of oil.
 16. The gas turbineengine as set forth in claim 15, wherein said engine is rated greaterthan or equal to 15,000 and less than or equal to 35,000 lbs in ratedthrust at take-off.
 17. The gas turbine engine as set forth in claim 12,wherein said oil tank holds greater than or equal to 25 and less than orequal to 35 quarts of oil, and said engine is rated greater than orequal to 15,000 and less than or equal to 35,000 lbs in rated thrust attake-off.
 18. The gas turbine engine as set forth in claim 17, whereinthe fan drive turbine includes an inlet, an outlet, and a fan driveturbine pressure ratio greater than 5, wherein the fan drive turbinepressure ratio is a ratio of a pressure measured prior to the inlet asrelated to a pressure at the outlet prior to any exhaust nozzle, and thefan rotor includes a plurality of fan blades, a fan pressure ratioacross the fan blades of less than 1.45, measured across the fan bladesalone.
 19. The gas turbine engine as set forth in claim 12, wherein thefan drive turbine includes an inlet, an outlet, and a fan drive turbinepressure ratio greater than 5, wherein the fan drive turbine pressureratio is a ratio of a pressure measured prior to the inlet as related toa pressure at the outlet prior to any exhaust nozzle, and the fan rotorincludes a plurality of fan blades, a fan pressure ratio across the fanblades of less than 1.45, measured across the fan blades alone.
 20. Thegas turbine engine as set forth in claim 19, wherein the fan blades havea fan tip speed of less than 1150 ft/second, and the gear reduction is aplanetary gear system.
 21. The gas turbine engine as set forth in claim12, wherein an oil capture gutter surrounds said gear reduction, andsaid gear reduction is a planetary gear system.
 22. A method ofdesigning a gas turbine engine comprising: providing a fan drive turbinedriving a gear reduction, said gear reduction driving a fan rotor;providing a lubrication system supplying oil to said gear reduction,wherein the oil remains in the oil tank for a dwell time of less than orequal to five seconds during operation, and said gear reductioncomprises a gear ratio greater than or equal to 2.3:1 and includes a sungear that drives intermediate gears; and said fan rotor delivering airinto a bypass duct as propulsion air and into a core engine where itpasses into a compressor section, with a bypass ratio defined as theratio of air delivered into said bypass duct compared to the volume ofair delivered into said core engine, said bypass ratio greater than orequal to 10.0.
 23. The method as set forth in claim 22, wherein saiddwell time is less than or equal to 3.0 seconds during operation. 24.The method as set forth in claim 23, further comprising oil bafflescircumferentially between said intermediate gears.
 25. The method as setforth in claim 24, further comprising an oil capture gutter around saidgear reduction.
 26. The method as set forth in claim 25, wherein saidoil tank holds greater than or equal to 35 and less than or equal to 50quarts of oil, and said oil tank is associated with an engine havinggreater than or equal to 35,000 and less than or equal to 100,000 lbs inrated thrust at take-off.
 27. The method as set forth in claim 26,wherein the fan drive turbine includes an inlet, an outlet, and a fandrive turbine pressure ratio greater than 5, wherein the fan driveturbine pressure ratio is a ratio of a pressure measured prior to theinlet as related to a pressure at the outlet prior to any exhaustnozzle.
 28. The method as set forth in claim 27, wherein the fan rotorincludes a plurality of fan blades, a fan pressure ratio across the fanblades of less than 1.45, measured across the fan blades alone.
 29. Themethod as set forth in claim 28, wherein the fan blades have a fan tipspeed of less than 1150 ft/second.
 30. The method as set forth in claim29, further comprising a mid-turbine frame positioned intermediate thefan drive turbine and a second turbine.